In operation of a gas turbine engine, air at atmospheric pressure is initially compressed by a compressor and delivered to a combustion stage. In the combustion stage, heat is added to the air leaving the compressor by adding fuel to the air and burning it. The gas flow resulting from combustion of fuel in the combustion stage then expands through a nozzle which directs the hot gas to a turbine blade, delivering up some of its energy to drive the turbine and produce mechanical power.
In order to increase efficiency, the nozzle has a preestablished aerodynamic contour. The axial turbine consists of one or more stages, each employing one row of stationary nozzle guide vanes and one row of moving blades mounted on a turbine disc. The aerodynamically designed nozzle guide vanes direct the gas against the turbine blades producing a driving torque and thereby transferring kinetic energy to the blades.
The gas flow typically entering through the nozzle is directed to the turbine at an entry temperature from 1850 degrees to at least 2200 degrees Fahrenheit. Since the efficiency and work output of the turbine engine are related to the entry temperature of the incoming gases, there is a trend in gas turbine engine technology to increase the gas temperature. A consequence of this is that the range of temperature gradients between starting, operating and stopping increases and the relative expansion between the sealing ring and the turbine blade must be compensated therefor.
Additionally, a seal ring is positioned about the turbine disc and moving blades. The seal ring thermally expands and contracts depending on the current temperature. To compensate for the relative thermal expansion, such as starting and stopping, a clearance or space is typically spaced between the tip of the turbine blade or rotor seal and the stationary seal ring. The larger the clearance between the seal ring and the turbine blades, the less efficient the combination. However, if the clearance between the turbine blades and the seal ring is not maintained, the blade tips will rub and interfere with the seal ring causing structural failure of the components.
Historically, nozzle guide vanes, turbine blades and sealing rings have been made of metals such as high temperature steels and, more recently, nickel alloys. In most of these applications, it has been found necessary to provide internal cooling passages in order to prevent melting. The use of ceramic coatings can enhance the heat resistance of nozzle guide vanes and blades. In specialized applications, nozzle guide vanes and blades are being made entirely of ceramic, thus, imparting resistance to even higher gas flow entry temperatures.
Ceramic materials are superior to metals in the high-temperature application due to a low linear thermal expansion coefficient.
When a ceramic structure is used to replace a metallic part or is combined with a metallic one, it is necessary to avoid excessive thermal stress generated by uneven temperature distribution or the difference between their linear thermal expansion coefficients. The ceramic's different chemical composition, physical properties and coefficient of thermal expansion to that of a metallic supporting structure result in undesirable stresses. A large portion of the undesirable stresses is thermal stress. Thermal stress will be set up either in the nozzle guide vanes, blades and/or sealing ring and between the nozzle guide vanes, blades and/or sealing ring and their supports when the engine is operating.
Furthermore, conventional nozzle, blade and sealing ring designs which are made from metallic material are capable of absorbing or resisting these thermal stresses. The chemical composition of ceramic nozzles, blades and sealing rings do not have the characteristics to absorb or resist high thermal stresses, which are tensile in nature.
The present invention is directed to overcome one or more of the problems as set forth above.